text:
Can someone identify if this is a Boeing or an Airbus?
I would like to know if this is a Boeing or an Airbus. Maybe there is an Aviation Wizard out there. It's an EasyJet plane.

I hope this is Ducktape (Panzertape).

I've just joined to ask; why are you concerned about the plane manufacturer and not "what the hell are they doing taping up the engine?"

an even more important question is "what is that pile of loose, silver colored stuff sitting inside the nacelle, and will our intrepid duct-taper clean it out before FOD?"

Is it just me, or does that molten metal casting in the nacelle look like a cast of a human hand?

@Darren I'm just writing an article about "Why Airbus is better than Boeing." I found this picture and if it was a Boeing, I would have included it into the article for fun, saying their planes are held together by Duct-tape :)

@Darren I've been here for quite a while and I'm wondering the same thing... Noah, this would seem to be counter-evidence to your article title. :P

@FreeMan Maybe it's a T-1000?

@NoahFisher All aircraft will at sometime have speed tape (as used here on the outside) and duct tape (frequently used on the inside).  I have never worked on or flown an aircraft that doesn't have tape holding something together.  It's not a "bodge" job.  They are recognised materials for approved repairs and are not an indication of either the quality of the aircraft or the quality of the maintenance.

There's no need to be an aviation wizard - Easyjet only fly Airbus A319 and A320 aircraft.
Even if you didn't give the airline, I would guess that it is an A320. Comparing the size of the engine to the person suggests it is a narrow-body aircraft, and 737 engines have a distinctive, flattened lower surface, whereas the A320 series are more cylindrical.

Amazing! I already thought about looking if they only fly a specific Company, but I thought: Nahh. When looking at your second paragraph: Thats what I meant by "aviation wizard". Thank you!

While it is an EasyJet A319/A320, its not as easy as saying "its EasyJet so it can only be an A319/A320" because EasyJet flew 737s until 2011, so you have to rely on engine shape rather just airline...

Sub-question: doesn't the small triangular intake on top of the engine also give it away? I've never seen it on 737s.

@peppe that "small triangular intake" on top of the engine is a NACA duct. NACA DUCT ALL THE THINGS!!!!  :)

Note that the newest 737s no longer have the familiar "hamster-pouch" intakes.

According to Airfleets.net, EasyJet has a full Airbus fleet. This engine colouring indeed seems to belong to EasyJet so, yes that is an Airbus A320 family.

Sure, it looks like EasyJet but it isn't sure to be. And EasyJet used to fly 737s, too. So "It looks orange and the only airline I can think of that flies planes with orange engines is EasyJet, who currently only fly Airbus planes" isn't very conclusive.

Although not definitive, this news story is for exactly this photo.  The Independent is not known for large inaccuracies or fabrications, unlike several other papers which also carried the same story so it's probably accurate in at least the basic fact of it's easyJet and they were taping the cowling.

@DavidRicherby: EasyJet had to explain publicly using speed duct is a common practice ("However an easyJet spokesperson told MailOnline Travel that in this instance, it was not used for repair work.") after the image made a hit on the net.

Thanks, Simon and mins. That's much more convincing than the unstated reasoning of the answer.

One could also look at the pylon which is very different on a 737 and doesn't have the NACA duct.


text:
Does the airflow from a propeller over a wing contribute to lift?
How I understand it, the perfect propeller would move air rearward at the exact same velocity as the free stream, i.e. the propeller air will merge with the free stream and have no real influence on the flow-field over the wing.
However a non-ideal propeller (thus all real propellers) accelerates the air to a higher speed than the free-stream and thus alters the flow-field over the wing.
My question is: Does this airflow from the propeller increase the lift generated by the wing?
My question obviously assumes that the propeller and wing is in the same plane and that the air from the propeller thus hits the wing.

Thrust is just lift by another name.

Except for a 90-degree mismatch in direction, perhaps?

A propeller not affecting the velocity of the free stream air would not generate any thrust (or drag), as the propeller works by the principle of creating thrust by accelerating a large mass of air, $F = m * a$.
The accelerated airflow from the propeller does generally increase the lift generated by the wing. The amount of lift increase depends on the configuration (how exposed the wing is to the slipstream of the propeller). This tends to be particularily noticeable on light twins:

(Image attribution: Wikimedia Commons)
There's also another effect on lift, coming from the corkscrew-shaped vortex of the propeller slipstream, perhaps more noticeable on single engine aircraft:

(Image attribution: FAA Pilots Handbook of Aeronautical Knowledge)
The image is a little bit unfortunate, as it demonstrates another effect as well (yaw from the slipstream), but you can also see how propeller slipstream will affect the flow pattern around the wing roots: on one side, it will tend to increase the effective angle of attack, leading to higher lift on that side, and on the other side it will decrease the effective angle of attack, decreasing lift on that side, leading to a roll-moment. This effect of course exists in twin engine aircraft too, but may be less noticeable, as the effect is more symmetrical, in the case of co-rotating propellers, or even symmetrical (in the case of counter-rotating propellers).

source: Stack Exchange

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sort: votes


text:
Are modern airliners still stressed-skin aircraft?
Can modern airliners still be seen as stressed-skin aircraft?
i.e. is the skin, whether it be aluminium or composites, still an integral structural member of the airplane?

Yes, absolutely. The skin transfers shear stresses between the stringers or longerons which run lengthwise along the fuselage and transmit tensile or compressive stress. Cutouts and openings (for doors and windows) need special local reinforcements, so the reduced structure can still withstand the loads.
Equally, the wing skin is the main element for the transmission of torsional loads. In today's airliners, the spar is part of the skin, but the wing skin ahead and aft of it is also a part of the primary structure.
Generally, stress follows stiffness. When a structure is loaded, it will deform, and those parts which withstand deformation most will carry most of the load. In the end, all parts which have similar stiffness will carry loads in proportion to their cross section. Since most of an airliner's structure is made from aluminum, it has the same stiffness.
See this question for one example clearly showing fuselage skin deformation.

Do you know if this holds true for the carbon-composite skin of the 787?

@raptortech97: That should probably be asked in another question.

@raptortech97: Yes, absolutely! Carbon-composites have similar to higher stiffness than aluminium (depends on fiber orientation) and show elastic deformation up to their breaking point, so their load-carrying capacity can and is fully being used. See this answer for more.

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text:
What is the difference between Boeing 73G, 73H and 737?
I found a data source, where Boeing 737 could be encoded with the following values - 733, 734, 735, 738, 73G, and 73H; Boeing 767 with values 762, 763; Boeing 777 - 772, 773, 777.
Any idea why such coding is used? Is it some standard to encode them with such values? Can I get more details from such codes (for ex., full aircraft type incl. customer code like Boeing 737-808)?

Where/what is "a data source"?

There are two main systems of aircraft type designations: IATA and ICAO codes. The ones you mention are IATA codes (all 3-character) and are often used for shortened aircraft type designation in public information.
Both code types are listed here with the corresponding full aircraft types. As you can see, the IATA information covers the type and variant, including some features such as winglets, but not necessarily extended range models or sub-variants with different engines. It also differentiates between freighters and passenger versions.
Despite the additional digit, ICAO codes are less precise and often include several variants such as freighter as well as passenger variants of the same type.
The codes given in your question translate to:

733  - Boeing 737-300
734  - Boeing 737-400
735  - Boeing 737-500
738  - Boeing 737-800
73G  - Boeing 737-700 (7th letter of the alphabet)
73H  - Boeing 737-800 with winglets (8th letter of the alphabet)
762  - Boeing 767-200
763  - Boeing 767-300
772  - Boeing 777-200
773  - Boeing 777-300
777  - ambiguous 777 type

(all passenger aircraft)
Unfortunately, as @Ralph J mentions, it's not always clear which type of code is used. Mostly, a 4-digit-code will be an ICAO code and a 3-character-code will use the IATA system, but because some ICAO codes only contain 3 digits, this isn't always reliable. For example, if you read "DC6", it could be the IATA code and refer to the passenger version, or it could be the ICAO code and you wouldn't know if it refers to the passenger or the freighter version.

The "733" is the 737-300.
The "734" is the 737-400.
The "735" is the 737-500.

Then there is the 737-700, which would, by that convention, translate to the "737".  But that's the general type, the 737 that encompasses all of the variants.  So you can foresee plenty of confusion if sometimes "737" means the Boeing 737, and other times "737" is short for the particular 737-700 variant.
So to avoid that confusion, they took the 7th letter of the alphabet, G, and used that instead of the "7" of "-700" and so now:

The "73G" is the 737-700.

For the sake of consistency, they did the some thing with the 737-800, and in some places you'll see that referred to as the "73H" -- as the 737-700 is followed by the 737-800, the 73G is followed by the 73H.  But in other contexts, you'll see the reference revert to the old convention, and

The "738" and "73H" refer to the 737-800.

The link in @JulianHzg's answer specifies which codes are used in IATA and which in ICAO usage, although sometimes you'll see a code in isolation without necessarily knowing which system of reference it's coming from.  In a pure, strict ICAO framework, the "B737" is ONLY the 737-700, with winglets, in a passenger configuration, but you'll see that used in places that aren't quite as strict in their conventions, and "73G" avoids confusion.

Interestingly, the 737-900 seems to always be "739" and never "73I".

Yeah, I'm a little surprised that "73H" gets used at all.  "73G" is the one place where you really NEED a letter to avoid the confusion "737" as "all" 737's, vs the 737-700.  There's no great risk of confusion over 738 or 739, so "73H" doesn't add any obvious value.  I'd suspect that "73I" carries the risk of being confused with "731" -- and even though that doesn't exist (very, very few 737-100's were ever built, and I don't think any of them still fly), it's a moment of confusion, so everybody uses "739" and it's perfectly clear.

73H has wing tips. 738 in the same source presumably does not, but other sources may refer to both as 738

While the existing answers cover the most common uses of such codes (ICAO and IATA codes,) in order to further complicate things, there is actually a third possible meaning, as well. Airlines often assign a letter designator in order to specify a particular cabin configuration. For example, Delta has no less than 9 unique codes for its 757-200 cabin layouts, even though they're all exactly the same airframe. If you're looking at a website like SeatGuru (or the website of the airline itself,) these are the codes you'll see in reference to exactly which cabin configuration will be on the particular flight you're booking.
A given seat number may refer to a completely different seat or even not exist at all from one cabin configuration to another, even on the same type of airframe on the same airline, so these codes are used to disambiguate the configuration. Unfortunately, this also has the result of making the code itself even more ambiguous in the absence of context indicating which type of code it is.
While this type of code doesn't mean much to the pilot, it can mean a lot to the passengers. To go back to the Delta 757-200 example, on the 75A configuration, seat 2A is a normal First Class seat with 40" seat pitch. On the 75S configuration, on the other hand, it's a 76-inch-long flat-bed seat. Slight difference there. While the differences in economy usually aren't quite as dramatic, they can still be rather large. For instance, on the 75S configuration, seat 19F has a lot of extra leg room due to a missing seat in front because of an exit door, whereas it's just a normal economy seat on the 75A configuration.
As far as I know, this code is distinct from the customer codes discussed in this question, which would instead take the form of 757-232 for a 757-200 delivered to Delta Airlines. The code discussed in this answer denotes the current cabin configuration of the aircraft and takes the same form as the ICAO/IATA codes, whereas the code discussed in the other question denotes the customer who originally ordered the aircraft from the manufacturer. So, while all 757-200s initially delivered to Delta Airlines would be 757-232s, they can (and do) have many different cabin configurations, each with their own code of the form 75A, 75S, 75Z, etc.

Thanks. This part of the question we discussed some time ago - http://aviation.stackexchange.com/questions/8379/what-does-3q8-mean-in-the-aircraft-model-boeing-767-3q8

@LA_ Actually, as far as I know, that's yet another code (and has a different form.) See the last paragraph that I just added.

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text:
What would happen when the fuel tank of a Cessna 172 became (partial) vacuum during horizontal flight?
Let's say we are flying a piston engine high wing airplane (Cessna 172) that uses a carburetor.
What would happen when the fuel tank of a Cessna 172 became (partial) vacuum during horizontal flight?

C172 has two tanks. Are you asking what happens if only one stops flowing?  Probably depends on  model year also, as I think the venting methods changed over the model years.

I don't understand your question. Can you tell us exactly what it means if a fuel tank "becomes vacuum"?

By vaccum, do you mean just empty or even without air (in depression)?

By becoming vacuum I refer to the pressure inside the tank that drops below atmospherical pressuure.

The 172S POH, the current version in production indicates:
FUEL VENTING
Fuel system venting is essential to system operation. Blockage
of the system will result in decreasing fuel flow and eventual engine
stoppage. Venting is accomplished by an interconnecting line from
the right fuel tank to the left tank. The left fuel tank is vented
overboard through a vent line, equipped with a check valve, which
protrudes from the bottom surface of the left wing near the wing
strut. Both fuel filler caps are also vented.
FUEL SELECTOR VALVE
The fuel selector valve should be in the BOTH position for
takeoff, climb, landing, and maneuvers that involve prolonged slips
or skids of more than 30 seconds. Operation from either LEFT or
RIGHT tank is reserved for cruising flight.
NOTE
When the fuel selector valve handle is in the BOTH position
in cruising flight, unequal fuel flow from each tank may
occur if the wings are not maintained exactly level.
Resulting wing heaviness can be alleviated gradually by
turning the selector valve handle to the tank in the "heavy"
wing. It is not practical to measure the time required to
consume all of the fuel in one tank, and, after switching to
the opposite tank, expect an equal duration from the
remaining fuel. The airspace in both fuel tanks is
interconnected by a vent line and, therefore, some sloshing
of fuel between tanks can be expected when the tanks are
nearly full and the wings are not level.

So if the pressure inside the tank drops below atmosferical pressure then fuel may not reach the carburetor?

Yes.  Take a jug of water with a small hole in the cap and tip it upside down.  Water stops flowing as the suction is created in the top.  The 150/172 didn't originally have vented caps.  They were were added later (under an AD IIRC) to provide a back up venting method if a wasp fills your vent line with mud. So the vent tube itself is kind of redundant.

@John K That sounds like a struggle for the last piece of ketchup with a squeeze bottle.

Except: fuel flows much more readily then ketchup, and fuel tanks aren't squeezable.

Interestingly, I saw recently on a youtube doc on ketchup that it has a unique physical property where its viscosity drops sharply when subjected to shock of a certain severity, which is why you have to whack the bottom just so to get it out.

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text:
What does the student pilot have to do during his/her PPL practical course?
The PPL theory consists of:

Principles of flight.
Flight performance and planning.
Aircraft general knowledge.
Meteorology, Navigation, Flight Planning and Monitoring
VFR Communications
Human Performance and Limitations
Air Law

I was wondering, what does the student pilot have to do during his/her PPL practical course?
It would be nice to see a list of things that needs to be mastered. A link to a practical resource would also help.

Please tag your question with a governing region (FAA, EASA, etc) so that we can give a more specific answer.

This depends on the jurisdiction/governing body that is issuing the license but for the FAA you can find a full list of everything covered on the Private Pilot Check Ride here your training should cover those topics and maneuvers. How they are covered is up to your instructor.

For the UK, the EASA PPL requirement for SEP(A) are listed here

If you don’t already hold a licence, you will need to complete 45
  hours of flight instruction on aeroplanes, 5 hours of which may have
  been completed in an approved flight simulator (an FSTD – flight
  simulation training device), including at least:

25 hours of dual flight instruction,
10 hours of supervised solo flight time, including at least 5 hours of solo  cross country flight time with at least 1 cross country flight of at least 270 km (150 NM) that includes full stop landings at 2 aerodromes different from the departure aerodrome.

You will need to complete a skill test after you have completed all
  the flying requirements.
You will also need to ensure that the examiner meets the requirements
  for test notification and designation as set out in:
IN 2014/110 (for UK examiners)
IN 2015/050 (for non-UK examiners)

The list on the left of that page also lists requirements for other EASA licence types.
Non-EASA aricraft can be flown under an NPPL in the UK.  An outline of the NPPL is listed here

You can also look at the UK CAA Standards Document 19, LAPL and PPL Skill Test (Aeroplanes), Policy and Guidance for Applicants and Examiners, which list each item that you need to perform individually, I used the Wings Taking Flight app that links to the document and other similar ones. For example:

(c) Climbing:
• Maintain directional control and balance throughout.
• Trim for nominated speed including best Rate of Climb speed (VY).
• Complete all necessary climb checks.
• Turn onto given headings maintaining balance and speed and bank
angle.
• Maintain lookout throughout.
• Return aircraft to straight and level flight in cruise configuration
at nominated level/ altitude.
• Complete all necessary drills and checks.
• Best angle of climb.
• Maintain heading and balance during transition from cruise or
descent at VSO + 10 kts to best Angle of Climb speed (VX).
• Complete all necessary climb checks.
• Turn onto given headings maintaining balance and speed and bank
angle.
• Maintain lookout throughout.
• Return aircraft to straight and level flight in cruise configuration
at nominated level/ altitude.
• Complete all necessary drills and checks.


text:
Are cargo and passenger flights flown differently?
Are aircraft flown differently because boxes don't complain about more abrupt changes in direction, speed, altitude, etc where passengers might. i.e. if the cargo is incapable of loading itself, is it subjected to a more aggressive flight regimen than if the cargo is self-loading and potentially complaint-filing.
I understand that Boeing, Airbus and other manufacturers produce essentially identical airframes in both passenger and cargo configurations, so they would have essentially identical flight envelopes, I just wonder if cargo flights are taken closer to the extremes.

Yes, cargo and passenger aircraft of the same basic model are sometimes flown differently when there is no live cargo. A friend of mine who flies cargo put it almost exactly as you did in your question when you say boxes cannot complain, so he sometimes hand flies the aircraft when passenger pilots would be required to use autopilot. Cargo pilots don't have to worry about tender stomachs in the back, so steeper climbs and descents and tighter turns are acceptable as long as the aircraft remains well within the flight envelope.

You can see the difference when comparing flightpath of the two directions of the Reykjavik-Liège weekly flight by Icelandair Cargo. Export is live horses, import is machinery. So, on which leg would the pilot choose a stomach-friendly flight path?

My friend does fly horses occasionally, dogs as well. With those flights he's very careful so he doesn't stress them.


text:
Why would a helicopter climb in a spiral?
I was just watching the James May's Toy Stories Christmas Special Flight Club*, and in it he commented that the helicopter was making a spiral climb to 8,000'.  Why would the pilot spiral, not just essentially go straight up?
* An excellent series if you're a big kid who enjoys toys, and as a bunch o' GA pilots, I'd think that covers most of you!

A helicopter will achieve its best climb rate at a moderate forward speed. Climbing in a spiral helps to have forward speed in what is essentially a vertical climb.
In a hover all the airflow which is available for lift creation must be generated by the rotation of the main rotor. This means that a small amount of air must be accelerated by a lot. If the helicopter adds forward speed, it can achieve a higher mass flow through the rotor, and now less acceleration is needed to achieve the same lift. This improves the efficiency of lift creation. If the helicopter goes faster than its speed for maximum rate of climb, aerodynamic drag grows too high and reduces efficiency again.
The difference can be dramatic: Below is a plot of the speed versus altitude of a generic helicopter and turboprop plane in comparison to the V-22 tilt rotor. Remember that the available power is roughly proportional to air density, and you can start to see how much more power is available for climbing when the helicopter flies with a moderate forward speed.

Flight envelopes of a generic helicopter, a turboprop plane and the V-22 tiltrotor (picture source). Thanks to @mins for inspiring me to expand the answer a little.

Does the "spiral" aspect of the climb mentioned above add anything to the efficiency, as opposed to climbing straight ahead?  It would seem like there would be a slight loss in efficiency since a little bit of the lift vector is being used to accomplish the turn.  (Obviously a much less pronounced loss at a ten degree bank than a 45 degree bank.)

@RalphJ: A straight ahead climb would probably be more efficient. But that means the heli isn't maintaining geographical position.

Would you think the same applies, with a quarter turn, to horizontal propellers? In other words would the efficiency of an aircraft propeller theoretically improve if some lateral motion was added to the forward motion of the aircraft? It is true that it does not seem practical, but would it in theory be more efficient if propellers were made say to rotate around the fuselage, making a spiral trajectory through the air?

@MarcvanLeeuwen sounds like an opportunity for a whole new question!

How large can the translational lift be (compared to the hovering lift)?

@mins: I would compare the maximum flight altitude to the maximum altitude in hover. Normally, the maximum altitude is maybe twice as high as the hover altitude. Look here for an example. Note that the lift is the same in both cases (the helicopter is the same, after all), but the power required for creating this lift goes down in proportion to the densities at the respective altitudes. So it can be said that the maximum lift can be higher since more power is available, but the increase is less than the density ratio.

Stunning! Thanks.

@mins.  I don't know, and will continue to look for emperical references, but translational lift is very siginficant.  It is possible to not have enough power to lift into a stable hover (hot day, full fuel, a couple of chunky chaps in the cockpit) but, you can creep along the ground.  At a few knots, around 7/8, the craft goes light on the skids/wheels.  At about 10-15 kts, you can then lift off and continue into a safe climb with a power margin (varies according to QNH and OAT).  In other words, you can only get airborne with translational lift assisting.

It's simulated and trained by the QFI saying "you cannot use more than xx inches manifold pressure or xx% torque.  Now get airborne".  It can take a lot of room - I've needed about 50 metres to do this sometimes.

See my question here.
The reason I performed a spiral climb is for the reasons given above plus a couple more.

Forward speed is safety.  In the event that the engine should quit, having forward speed gives you a lot more options.  When climbing vertically, you have a lot more pitch on the blades than in forward flight.  If the engine does quit, the blades will slow down very quickly reducing the time available for the pilot to dump the collective and enter autorotation.  You are also storing kinetic energy, as well as potential energy, in the system and therefore have more energy to play with if the donkey does go quiet.
In the R22 I was flying, I would probably have about one second to react correctly in a vertical climb.  Perhaps 2-2.5 in forward flight.
I wanted to remain clear of the danger area since I didn't have a clearance and had limited options for flying around the sides.  I did ensure that I was not over any houses and moved the spiral a couple of times so that my noise footprint was not over the same spot.

In the spiral, the helicopter generates translational lift, which is added to the lift generated by the downwash of the rotor system, improving its ability to climb.


text:
Is there a special ULD to be loaded at the main cargo door?
Watching this video, it appears that the hydraulic ram intrudes into the cargo space and would run into the Unit Load Device (ULD) that was stowed there.
Is there a special ULD that is loaded on each flight to clear the ram, or is there enough clearance that any ULD will fit in this spot?

No, the container near the door is the same as the others. The containers do not fill the entire cabin all the way around, which leaves enough space above for things like the door hydraulics. The containers on the upper cargo area below are all 88" x 125", which corresponds to a standard LD7 container.

Source


text:
What is this pod at the wing-tip of a U2?
In this image of a U2S, what it the very unaerodynamic looking pod at the wing-tip? It looks like it holds some sort of sensor package. However, it doesn't seem to appear on all U2s (at least not based on photos on Wiki).

Source: Wikipedia

I believe it's part of/an attach point for the landing skids. That may not be the only use, however, so I'll pass on leaving this as an answer

The pod contains the Radar Warning Reciever (RWR). The variant in the picture is the TR-1B, which had wingtip mounted RWR units. According to globalsecurity.org:

The Air Force has fielded new electronic countermeasures on the aircraft, giving it a jamming capability and upgraded radar warning equipment with wing-tip mounted digital receivers.

As an another exmaple, SAAB seems to have put RWR on the wingtips of Gripen, though the equipment are certainly different (U-2 uses AN/ALQ variants).
Note that the bottom of the 'pod' has (wingtip) landing skid at the bottom and navigation light on the side. The RWRs face outside at an angle of ~45$^{\circ}$. This (containing RWR, skid and nav lights in a single package) is a later design and was added to the exiting U-2Rs as retrofit. It was not avialble in the earlier models.

Your AN/ALQ link seems quite rather broken.

@FreeMan Fixed it now.

In here you can find the handbook of the "civil" version on the U2, the ER-2 by NASA. Researching here and there it seems to be a "pod". In particular from the document above:

Other payload areas that have been used in the past, including the
  wing tips, [...]  can  support  small  instrument  assemblies. The
  structural and electrical connections for these non-standard installa-
  tions require individual consideration. Contact Lockheed Engineer-
  ing  for  attachment  details  if  you  are  considering  one  of  these
  locations.

What is inside it depends on the versions.
According to this document in the 90s they were proposing the wing tip to install potassium magnetometers to make high altitude magnetometric measurements ove Alaska.
According to this site it is a pod used for Electronic Counter Measures (ECM).
The ECM has been upgraded in 2005, but from the photos on Airliners.net it seems to be there since before 2005.


text:
How does a pitot plate work?
I understand the basic concept of the pitot tube in that it measures the air pressure coming into the tube in order to compare it to a static pressure measured elsewhere on the aircraft.
The B-2 bomber, however, has pitot plates in order to eliminate the radar reflectivity of the pitot tube.
How do these plates work, as compared to tubes, and why aren't they more common, since a flat surface that is less radar reflective would also, likely, be more aerodynamic?

A plate might possibly more aerodynamic, but the implied question is whether its materially more aerodynamic. Also, aerodynamic efficiency is only one criterion for the pitot system: other criteria are ease of certification, accuracy, manufacturing and sourcing difficulty, etc.  So I'm OK with the "how does it work" part of the question, but the "why aren't they more common" part needs to be justified.

 Pitot tubes are common, simple, cheap, understood by everyone.  Pitot plates are not common, complex, ludicrously expensive, understood by almost no-one.

Not sure why this got a down vote. Does anyone have any ideas on an answer to the question?

Something interesting here... http://www.pbase.com/garyhall/image/105738953

The B-2's pitot plate has actually been the cause of an accident in conditions that may or may not have had the same effect on aircraft with pitot tubes.
Pitot plates work much like pitot tubes; they measure the pressure of incoming air. The main difference is just how they measure it: pitot tubes measure the ram force of air while pitot plates measuring how much air is constricted by an obstruction. Pitot plates direct air through a "plate" with a hole that impedes the flow almost like how a Venturi tube accelerates fluids. Pictured: Pitot/Orifice plate

Pitot tubes are easier to maintain, inspect, and install. Some of their main problems that you might not encounter with a pitot plate, such as icing, are countered with tested and approved safety systems (like heating). In addition, pitot tubes are not the only instrument to measure airspeed; most commercial and GA aircraft use pitot-static systems in which other pressure sensors are used to get a better and more reliable reading. Pitot plates are advantageous in that they reduce air pressure about 5:1 rather than the pitot tube's 10:1 (meaning yes, less drag in some sense) but on such a tiny scale it makes barely any difference. On the flip side, pitot plates have very inaccurate readings at lower speeds (though so low that it probably would not matter for any jet aircraft). Also, it is to my knowledge that pitot plates are used mainly for fluid research, sometimes not even related to aviation.
Since pitot tubes are relatively small compared to the size of an aircraft and are streamlined anyway, there would be little to no noticeable improvement in aerodynamics. You would think that jetliners have completely clean surfaces, but guess again:

Removing a pitot tube won't make a difference. Note: the gray circles outlined in dotted red boxes are pitot-static ports, which, combined with an aircraft's pitot tubes, make up the pitot-static system. They can measure air pressure for both airspeed and altitude among other things.
Final thing I wanted to clear up: flat surfaces are generally less aerodynamic than curved surfaces. Someone got too much of that idea when designing this plane... /s


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Why is a higher bypass ratio more efficient at lower speeds?
Why is it that a larger volume of air moving slower is a more efficient means of propulsion than a smaller volume of air moving faster?  (This is true up to a given speed, obviously, as supersonic aircraft use lower BPR engines).

FTR, see this (my) related question on physics: Turbofan: Efficiency wrt. the bypass ratio: “a lot of slow air > a little faster air”?

Because bigger reaction mass with lower speed carries away less energy for the same momentum.
To produce thrust the engine applies force to the exhaust gasses and air. This force accelerates the air. To produce thrust $F$, the engine has to impart momentum $p = \frac{F}{t}$ per unit of time on the air. Momentum is proportional to mass and first order of speed. $p = mv$ where $m$ is mass of the gasses (reaction mass) affected by the engine per unit of time and $v$ is speed of those gasses.
The gasses however also take away some kinetic energy. Kinetic energy is also proportional to mass, but also to second order of speed. $E_k = \frac12mv^2$.
This energy has to be provided by the engine. Since the fuel can only provide so much energy (it's heating value) per unit of mass, an engine that needs less energy needs less fuel. And that is achieved by moving more air. Which means higher bypass ratio for jet and bigger propeller for engines with propellers.
Note, that this does not say high-bypass engine would be more efficient at slow speed than at high speed. Because it isn't. All jet engines are more efficient at high speed. This is because the power the engine produces is $P = Fv$. So at high speed, larger part of the power output is the useful work moving the aircraft.
In high-bypass engine the maximum power output of the fan is limited though (similar to propeller), so compared to a turbojet the difference in efficiency is higher at low speed.

Interesting parallel: The same goes for lift: It is more efficient to produce the same lift by moving a lot of air by a little. Like a high bypass ratio engine, a high aspect ratio wing is more efficient. The physics is the same in both cases.

@PeterKämpf won't you mention wing loading?

@Abdullah Why should I? The parallel here is that both have a quadratic dependency over speed so there must be an optimum somewhere. Wing loading has nothing to do with this.


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What advantages are provided by an unstable canard configuration for fighters?
Why are most canard fighter aircraft unstable?  A stable design would let the canard provide lift.  My guesses are:

Increase pitch rate
Decrease stability while supersonic (supersonic, the aircraft are stable)

You want your control surface to create only little lift, so it has enough margin for maneuvering. This confirms your first guess.
Your second guess is also correct: Flying supersonically shifts the center of pressure back and increases stability. If the aircraft is unstable subsonically, it will suffer far lower trim losses supersonically.
Please read all about unstable designs in this answer. For conventional, stable configurations it makes most sense to put the control and damping surfaces at the back, and a canard is the better choice for unstable, highly maneuverable aircraft.

A canard surface operates in 'clean' air, not 'shadowed' by the wing, so maneuverability has priority over lift.


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Can the thrust of a turbofan be increased by increasing bypass ratio?
I read that the CFM56 is based upon the F101.  Id on't know how true that is, but it generated the following question:  If a larger fan is fitted to a low bypass turbofan, would dry thrust be increased?  Or is it the same amount of thrust produced with less fuel consumption?  That is what I was told on another message board but it doesn't seem to make sense.  Assuming that the proper size turbine is fitted, the same turbine airflow should be turned into more thrust with a larger fan.  Which is correct?

I liked the question, it made me think.

Non-afterburning thrust

Fitting a larger fan does two things to the thrust of an engine:

The static thrust is increased. Static thrust is the thrust produced when the engine is not moving.
The thrust gradient over speed becomes more negative, which means that thrust will decline more with speed when the bypass ratio is increased.

But there is more: Exit speeds at the nozzle are reduced, and a bigger intake is required, since the bigger fan will need more air. Increasing the bypass ratio means taking some of the kinetic energy of the core flow and turning this into a higher fan mass flow.
If we look at the formula for the propulsive efficiency $\eta_p$ of an airbreathing engine:
$$\eta_p = \frac{v_{\infty}}{v_{\infty} + \frac{\Delta v}{2}}$$ where $v_{\infty}$ is the speed of the engine and $\Delta v$ the speed increase of the gas flowing through the engine, the speed dependency becomes clear: When $v_{\infty}$ is low, a smaller $\Delta v$ acting on a higher mass flow makes the engine more efficient. When $v_{\infty}$ is high, however, this effect disappears, and now the smaller engine with a smaller and lighter intake becomes more attractive.
When the core of the engine stays the same, its fuel consumption will also stay the same, but the bigger fan will create more thrust, especially at low speed. The core mass flow will be the same, regardless of fan size, and the amount of fuel to heat this mass flow will also not change.
Since efficiency is defined as thrust per unit of fuel consumed, the bigger fan will also increase efficiency.

GE used the core of the F101 engine as the core of the CFM56 (with some modifications, presumably). The core of a turbofan consists of the high pressure compressor, the combustion chamber and the high pressure turbine. In this case, we talk about a 9-stage HPC and a single-stage HPT. Everything else was newly developed for the high-bypass turbofan.
Unlike just putting a larger fan on the low-pressure shaft, this is possible becasue the high-pressure components and the low-pressure components are relatively independent in a turbofan, i.e. the core of the engine is not affected too much by the size of the fan, the bypass ratio or other changes in the low-pressure area.
The new low pressure turbine (4-stage instead of 2-stage) can operate at a lower rpm and transmit more of the exhaust gas energy from the engine core to the larger fan (the exhaust gas coming out of the new LPT is probably cooler and slower than that coming out of the F101 LPT). The new low-pressure compressor increases the overall pressure ratio and therefore increases the amount of energy that goes to the low pressure compressor and fan, thereby improving the overall efficiency. This is probably possible because the non-supersonic operational envelope of the CFM56 means that intake air total temperatures will be much lower so that the same EGT margin can be achieved with a higher pressure ratio. In other words, becasue the incoming air is cooler, it can be compressed further without damaging the high-pressure turbine because it is too hot.
The static thrust of the CFM56 is 19500 lbf, which is up about 15 % from the F101's 17000 lbf without the afterburner. While a small part of this is probably an aerodynamic loss due to the unused afterburner chamber, most of it has to come from the better use of the energy offered by the engine core exhaust in the larger fan. The difference might even be higher becasue the maximum thrust of the F101 might be adjusted for a much lower mean time between shop visits. I'm not sure about that though.
This shows that the main benefit of the high-bypass low pressure section of the engine is a significantly increased fuel efficiency, an increase in the static thrust and (much) lower noise levels. The F101, on the other hand, can operate at higher speeds, produce more thrust at high speeds, is lighter and has a much smaller cross-section.
(Edited per comment by fooot, thanks for finding the source.)

This page says 17,000 lb military thrust, which is a little lower than the CFM56.

The jet engine propulses the aircraft based on moving air behind the aircraft at a higher speed than the airplane. Turbofan tries to move a higher quantity of air at lower speed that is more efficient. That is what is called propulsion efficiency.
This can be easily understood thinking about how can you propulse yourself when you are in a skateboard, will be more efficient if you use your hands over a wall or over something that can slip on floor? The heavier that "thing" the best for your propulsion reaching the building wall.
Ok, that's a good concept, but you need to do it!!! For being able to do that you need to connect a turbine at the exhaust of the previous turbine connected mechanically to the new fan of the engine.
Easy but no so good...

You are adding a new compressor stage (the fan) in front of the previous stages. So you need to make sure that all your materials are prepared for it.
You exhaust of the core might not be properly prepared to the new conditions.
Probably external shape of the previous engine is not prepared for a secondary flow.
Systems need to be reviewed.

In principle, with some modifications, that operation you suggest will work but just if we are talking about small modifications and not going to significant bypass ratio, where modifications will be of the same cost of a complete new engine design.
Also I see economical difficulties, in the current context there is significant competition for making the lowest comsuption. If you perform what you are proposing you will have an engine optimized for not having secondary flow adapted for secondary flow competing with engines optimized for working with secondary flow. Engine will be unefficient compared with competitors.
Finally, you will increase weight and thrust.


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Why would different crossing runways at one airport be on different radio frequencies?
In this answer the following image was included which shows that control of two crossing runways are on different radio frequencies.

Source: AIP Denmark EK AD 2 - EKCH - AREA OF RESPONSIBILITY
It seems to me that having crossing runways on different tower frequencies could lead to a runway incursion or worse, an accident, if there is any loss of controller-to-controller communication within the tower. One of the key tenants for pilots is situational awareness, and being on a different frequency means that the pilot of an aircraft receiving clearance for take-off on 12 won't know about a plane receiving clearance for take off on 22L at the same time and thus, couldn't question the clearance.
Now, I realize that it's unlikely for both 12 and 22L to be in use at the same time, but if the winds are fairly calm and to the North, it could be possible, especially if traffic was high. However, I seem to recall some discussion here about an (near?) accident that happened on crossing runways where the pilots were unable to see each other as they began their their take-off rolls.
My question is: Why would different, crossing runways be assigned to different tower frequencies since this can lead to a lack of situational awareness and potentially cause incidents?

There are many reasons that airports might imply different frequencies, including:
Frequency congestion (more frequencies, less aircraft on a frequency, less congestion),
Ease of controllers (It's easier for the controllers to control smaller parts).


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How much distance do I need between quadcopter propellers to avoid issues?
For a given propeller size (say 5"), what is the minimum clearance needed between propellers to ensure that not only do they not hit each other (easy to calculate), they also don't interfere too much with each other (i.e. tip vortices, airflow, etc.)
For example consider the following design: an X-shaped quad-copter where the propellers are at the end of each tip of the X - and the X across the diagonal (max distance from the center of the propellers) is 180mm - then 5" propellers will only have .279mm of clearance between them - now assuming that mechanically there are no vibrations/imperfections/etc. that cause them to run into each other, what else is there to avoid in terms of airflow? if they were raised such that two diagonal propellers are lower/higher than the others, would there be any issues? does it depend on the propeller pitch, or if it's a 2-blade or 3-blade propeller?

P.S. - I am unsure of the tags, please adjust as necessary

There are forums like rcgroups specialized in quad-copters (see:   http://www.rcgroups.com/aircraft-electric-multirotors-790 ).

I'm voting to close this question as off-topic (and too broad). It belongs to a RC site, as already pointed out.

@Energizer777 likewise there are also forums specialized in aviation - I did also try asking in the meta prior to posting and asking in the chat, so if it is indeed off-topic, I think that there is some sort of system flaw here and you should update the help center accordingly.

I agree that this would be better answered on a site more devoted to RC. But spacing of full-sized props would be on-topic here, and that is just a matter of scale.

I think this is a valid question for this site. The fact that another better forum might exist by itself does not seem reason enough to close it on here.

IANAAE: QC blades spin so fast that I simply cannot fathom any tip vortice issues at all, nor much, if any, interference between the rotor disks.

Head off to your local hobby/RC store, Best Buy, Walmart, etc and take a look at the commercially produced ones. Seems to me there are some pretty small ones available. Take a tape measure and get some measurements (if they have some out of the box) - that should give you some idea of where to start.

@FreeMan if all I wanted was measurements, I could just get those from online schematics - somehow as a "new" field, I don't see the local stores having very "well-engineered" designs, rather just "easy to get to work" or stuff from a more retail-oriented perspective (avoid returns) than sound design decisions. They are also horrible as far as which is faster, more agile, efficient, etc. as they will all claim all of those for marketing reasons and not be comparable across models.

what else is there to avoid in terms of airflow.... On-topic.


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Have contrails ever occurred at low altitudes, such as at takeoff?
Contrails form from the exhaust of jet engines condensing to water-ice. Has this ever happened at low altitudes, especially during or immediately after takeoff?

Related: How does contrail formation differ from turbofan to turbojet?. Have a look at the diagrams.

This is entirely possible if it's cold enough and has also happened. Here's a video of a Tu-204 taking off from Yakutsk at -45°C.

Source

That's pretty awesome, reminds me of a rocket plane. So much so in fact, I think it deserves to be the answer.

@DrZ214 Yeah, looks so cool

@ymb1 Nice find!

Photos of this would be very hard to come by, but yes, in an extremely cold environment, say close to the arctic and antarctic regions of the globe where daylight temperatures can be down around -40°F, it is entirely possible for this to happen by the laws of thermodynamics.
The rarity is further exacerbated by the fact that aircraft are difficult to start up at those temperatures, generally grounding operations where surface temperatures reach low enough to form contrails.

Contrails also form behind piston engines. I used to fly Cessna 402's in -40C conditions in the Canada and would often see white contrails behind each engine.  I also once saw a Cessna 150 doing circuits in -40C conditions and it too had a short contrail behind it.

I can believe it.

Here you go:
http://www.airliners.net/photo/Aeroflot-Russian-Airlines/Sukhoi-SSJ-100-95B-Superjet-100-%28RRJ-95B%29/4166253/L?qsp=eJx1jbEOgkAQRP9laxpEvYROGkst/IHN3qiXIFz2toAQ/t3jSOzsJu9NZhaScTBM9pgjqKUEVnlTRZGVP4nahTioKD%2Bt4xRkr9WHU%2BOqn7ligP5x1B5d7TJIo1o35wvPhosIosHTzm/qoZtCkvL92oZygN5LpuacuQ8p9lw2YBx6WtcvTfI/CA%3D%3D

@Thaumaturgic that may be contrails after all, but the lighting makes it appear like the poor plane has engine fire.

@MikeSowsun Did you ever see those contrails during or right after takeoff? Or was it just up high as "normal"?

They start at takeoff. They only extend a few feet behind the aircraft and imeadiately dissipate.

Not necessarily; it all depends upon the humidity of the atmosphere.  Contrails would remain for extended periods of time if the air was saturated.


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Are pitot systems protected from ice?
Aircraft use pitot tubes or pitot-static tubes. Do these have some kind of anti-ice system? If so, how? And if not, why not?
I'll just point out that it doesn't seem trivial to de-ice such a thing. A simple heating element at the front would melt the ice but then the water will travel through the tube and probably quickly reform somewhere else. So the entire tube could be heated, but then, that will drastically alter the ambient conditions that the pitot is trying to measure.

Pitot heat - https://aviation.stackexchange.com/questions/34557/why-not-have-pitot-heat-always-on

You can see an example sketch of a heated pitot tube in the FAA Handbook of Aeronautical Knowledge:

As you can see, there are two heaters that heat the entire tube and (ideally) stop ice from forming in the first place (anti-ice, not de-ice).

Thank you. My question should have been, how does such a heating not adversely affect the ambient conditions its trying to measure? But even that has been asked already. https://aviation.stackexchange.com/questions/34527/how-does-pitot-heat-influence-airspeed-indication-assuming-no-ice-or-moisture?rq=1

They also really need an anti-hornet nest feature :)


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Has there ever been a sideways H-tail on an airplane?
I've seen several aircraft with 2 vertical stabilizers and 2 rudders. This could be an H-tail like the B-25 Mitchell or the A-10 Warthog, or something more sleek like the Su-27.
However, I got to thinking, are there any airplanes that have 2 sets of horizontal stabilizers and thus 2 pairs of elevators? This would be a sideways H-tail. I guess it would be sort of like a bi-plane, but in the tail only.
Do these exist somewhere? Why or why not?
Experimental planes are okay.

Two small vertical stabilizers are sometimes used in place of 1 single big one, which is usually the highest point of an aircraft. This way you can lower the height of the plane and get a bigger plane that fits into small, old hangars. Using double horizontal stabilizer doesn't give that advantage, as the width of a plane is determined by it wings anyway, so you can make the tail pretty much as wide as you like, no need for doubling.

The double decker horizontal tail was a popular choice for large biplanes, even into the 1930s. Below is a picture of the Zeppelin Staaken R VI giant airplane and below it a Handley-Page H.P.42 passenger aircraft from 1931.

Zeppelin Staaken R VI over Leipzig (picture source)

Handley-Page HP42 passenger aircraft (picture source)
As with the wing, a braced biplane tail makes for a stiff, lightweight structure. An unbraced biplane tail was also chosen by Hermann Dorner for his light attack aircraft, not for weight reasons, but to reduce tail span in order to allow the tail gunner a larger firing arc. The picture below shows a Hannover Cl II. I guess this design comes closest to the requested "sideways H tail".

Hannover Cl II with raised tail, probably for compass calibration (picture source)
I have yet to find a monoplane with a biplane tail. Biplane tails vanished together with biplane wings.

Great finds, thank you.

Did the Hannover CL have elevators in all horizontal stabilizers? Or just the bottom ones?

@DrZ214: Both planes had elevators. See here for an overview drawing.

Most of the initial aircraft made by Wright brothers had the twin horizontal stabilizer setting. However, the elevators in these cases were set in the front, rather than back, as is the norm now.

By Wright brothers - OhioLINK Digital Media Center, Public Domain, Link
Quite a few of the early aircraft had a 'box' like configuration, notably the Santos-Dumont 14-bis, which made the first public powered flight. Again, though the pitch control was through the box like structure, it was more of a canard than a horizontal stabilizer.

Santos-Dumont 14-bis; image from Wikipedia
One of the earliest aircraft with a 'proper' twin horizontal (as well as vertical) stabilizer was the aptly named Bristol Boxkite

Bristol Boxkite; By Hpeterswald - Own work, CC BY-SA 3.0, Link
Another good example is its predecessor, the Farman III, which had two horizontal stabilizers and a single vertical tail right in the middle, forming a perfect sideways 'H'.

Farman III; By C.C. Pierce & Co. - http://digitallibrary.usc.edu/cdm/ref/collection/p15799coll65/id/12574, Public Domain, Link
You can see that most of these are early aircraft, which had structural limitations. As the aircraft design matured, the 'conventional' layout became the norm as a single stabilizer assembly can be used to achieve the necessary pitch control.

When the stabilizer is in the front, is it still a stabilizer? Or a destabilizer?

Also, the 14bis was not, strictly-speaking, an airplane (no full three-axis control).

There is quite a famous example of this!

but other than the Wright Flyer, I don't know of any.
My layman conjecture as to why this is not more common is that the tail only needs to be large enough to counteract the moment that CG exerts relative to the center of pressure of the wing. Large tail surfaces can balance large torque loads, but that is aerodynamically inefficient compared with maintaining a reasonable CG range.
Another aspect of that particular concern is that you can alternatively just make the tail boom longer to create a larger torque arm for the tail surface.
My answer to "why not" is that it's just not necessary. There are too many other ways to address control and stability issues to justify increasing the complexity with a double-horizontal stab.

Ha! Never noticed this about the Wright Flyer. Maybe I should specify modern control surfaces, and a vertical stabilizer.

Some truly magnificent photos of men in their flying machines!

The only recent example I can think of (and it's still only in the design & prototyping stage so may not even count) is the Beriev BE-2500 super-heavy amphibious craft.

This has a couple of very unique features; first and foremost, it's designed to operate either as an aircraft or as a ground-effect craft, flying very close to the ground on a cushion of air. This is why it has such an enormous wing area, and why the elevators have to be at the top of the rudders so as to not disturb the airflow around the wings. Secondly, it's also amphibious - with two large pontoons providing buoyancy and balance in the water. These are set as far apart as possible for balance, and are structurally connected to the rudders, giving a unique layout.

Straight out of Kerbal Space Program...


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Where is the generator in a large turbofan of a commercial airliner?

So, where do you put the electrical generator?
If it's gonna be coaxial, the only place is the rear (it will be burnt) or the front (it will get in the way of intake).
If it's gonna be coupled to a gear, I can't see how to mesh a gear onto the spinning parts, because the bypass air surrounds the entire engine.
Note: I'm interested in large turbofans of large airliners like the A380 or Boeing 787, basically anything with turbofans carrying 200+ passengers.
(P.S., does the generator spin from the LP turbine or the HP one?)

"bypass air surrounds the entire engine" - I think this is your misconception, there is a set of support struts and bearings which are responsible for keeping the spinning parts centered in the turbine.

@pjc50 I figured there had to be support struts somewhere, but figured they would be streamligned (very thin) and not able to support a generator. Now I see there is a radial gearshaft to drive some aux stuff, which to me seems very exotic and prompts more questions.

Image source
The generators are usually driven from the Accessory Gearbox. This CAD drawing is from the company that makes accessory gearboxes for the Rolls Royce Trent engines driving the A330 and A350, and it shows the connection of the accessory gearbox to the high pressure rotor.

This picture from the book The Jet Engine by Rolls Royce shows the same arrangement: a bevel gear drives off of the HP spool, then drives another bevel gear connected to the crescent shaped Accessory Gearbox underneath the engine.

The gearbox itself is shown above (from the same book), with the accessories connected to the LP spool:

The Engine Driven hydraulic Pump
Fuel pump
Oil pump
Pneumatic starter, requiring the highest loads and therefore dimensioning the gear strength.
The IDG: Integrated Drive Generator which provides the AC power.

The accessory gearbox contains the engine starter as well, pneumatic for older types and electrical for bleedless engines. The engines are started by turning the HP rotor, and the accessory gearbox must be connected to the HP rotor for this purpose. For this reason, the HP rotor drives the generator and other accessories, although it would be more beneficial for fuel economy to use the LP compressor. Source: The Jet Engine issue 5 by Rolls Royce.

IDG Location: Dedicated pad on the accessory gearbox

If it's gonna be coaxial, the only place is the rear (it will be
  burnt) or the front (it will get in the way of intake).

That's a good question and remark. Actually the generator (integrated drive generator or IDG) is off engine axis and goes on the corresponding IDG drive pad of the accessory gearbox (AGB).
The accessory gearbox is a gear train with multiple teethed wheels enclosed in a housing, located at the bottom of the fan case, or on its side when the engine nacelle would be too close to the ground. AGB for the CFM56-7B (Boeing 737 NG):

CFM56-7B AGB located on the side of the fan frame. Source.
Same engine, with accessories in place:

Accessories in place (NTSB inspection after fatal accident). Source.
The IDG is visible below the starter (the large cylinder with its long vertical bleed air pipe), identifiable by its 3 colored phase connections and wires also running vertically to the pylon connectors).

How is AGB driven by the engine?

Does the generator spin from the LP turbine or the HP one

HP (N2) shaft. I believe the first reason to do this is because the AGB is also used to start the engine, and the starter must drive the high pressure compressor rather than the booster.

If it's gonna be coupled to a gear...

Yes there are several gears in the way.
On the CFM56-7B engine, the N2 shaft speed is about 14,400 rpm, but the fuel pump works at 6,000 rpm, the IDG is more efficient at 8,000 and the hydraulic pump operates at 4,000. The purpose of the three gearboxes and the gear train in the AGB is to provide individual accessories with a support and a shaft at the desired rpm.
The wheels in the AGB are driven by the high pressure shaft (N2) using the inlet gearbox (IGB) located near the N2 shaft within the fan frame hub. Internal configuration, same engine:

How is the AGB driven by the engine high pressure shaft

The gearbox on N2 (inlet gearbox or IGB) rotates a radial shaft.
The radial shaft, which runs within one of the fan frame struts to cross both the core airflow and the fan airflow, drives a second gearbox located outside of the engine, attached to the fan frame case
The second right-angle gearbox (transfer gearbox or TGB) drives a horizontal shaft which itself drives the AGB gear train.
The AGB provides the drive pads with their shaft.

Same accessory gearbox with related components, ready to be driven by N2 shaft:

CFM56-7B AGB with its TGB and radial shaft. Source.
Allocation of dedicated pads on the AGB

Accessory pads on the CFM56-7B AGB. Source: CFM56-7B Engine Manual
AGB variations
Each engine has a different, but similar accessory gearbox. Another example:

AGB on a Rolls-Royce engine. Source.
Two electric generators actually and possibly the starter used as a generator
Note that on both AGB, there are two electric generators:

One is the smaller engine electrical generator (control alternator on the first AGB) which is used by the engine itself to feed its own electric equipment (notably the FADEC/EEC which controls the engine). This generator takes over the aicraft electrical supply (IDG) as soon as the engine is started (after the speed has reached 15% N1). The engine will continue to run in the event of aircraft electrical failure.
The other is the larger IDG, the main source of electric power for the aircraft.

The starter is another accessory. When starting the engine, the radial driving shaft is used in the other direction, the accessory box drives the engine shaft. Sometimes the starter and the generator are a single component, an electric motor being reversible. See How is a turbine engine rotated for starting and inspections?

You mean that itty-bitty thing...is the generator? Just how much power does it generate?

@DrZ214 compared with land based electricity supplies, aircraft supplies often run at much higher AC frequency (e.g. 400Hz not 50 or 60) which means the generator can spin faster. That reduces its size and weight, for a given power output. Nobody really cares how much the generators in a power station weigh, or how big they are, but weight is critical in plane design.

Yes, at the higher frequency, it has all the copper wire, but far less iron core.

@mins So you tell me that the handheld thingy in the photo generates over 100 kW?! That's incredible!

@yo': Yes, look at this one... 150 kVA :-)

It isn't so much frequency that allows the generator to be smaller, but its output voltage. The higher the voltage, the less current for the same power; the less current, the thinner the wires can be - less copper, so less mass and less volume. Higher frequency is beneficial because higher frequencies couple better in transformers, so they can be smaller and lighter too.

What is itty bitty, and what is hand held? I don't see any scale shown in the photos or referenced, and some of those engines are huge!

@yo' Where do you see handheld anything? In the photo from the Southwest uncontained failure, you can see that the diameter of the IDG (big black can that the wiring goes to) is roughly comparable to the length of the nearest guy's upper arm, so definitely >30cm. See the answer linked by mins for better photos.


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How do engine doors work?
"Doors" here refers to something in or around the engine that can close or open, stopping or allowing the flow of air (or some other gas).
I've heard of such things in a few engines. GE9X mentions "bypass doors between the booster and high-pressure compressor which open into the airflow path to reduce the chance of ice crystals entering the core." The J58 has many doors which control the overall airflow path in or outside the core.
How do these doors maintain a good seal in the closed position? Are they always rectangular? Do they need to be perfectly airtight, and if not, how good is good enough? What type of actuator is used (hydraulic? worm gear?)? How do they withstand the high heat?
I'm especially interested in doors near the turbine because those will suffer enormous temperatures. And as far as I know there are no o-ring materials that can work at those kinds of temperatures. How do they not get stuck in either position after sitting that way for a long time in the heat?

You typically have doors in the cowling forward of the compressor that are spring loaded closed and are pulled in by low pressure inside the inlet to allow more outside air into the compressor.
Inlet disruptions can cause incipient compressor stall, which creates rumbling sounds emanating from the engine.The suck-in doors around the cowl admit extra air when the compressor needs it. The rest of the time, the springs will hold the doors closed. They will be sealed with a rubber perimeter seal of some kind. They are sometimes called "De-rumble Doors".
The other type of doors are the ones used to reverse the airflow for reverse thrust on turbofans, called "Blocker Doors". They will be actuated directly by linkages connected to the translating cowl section in the case of a cascade reverser.


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What is the speed of gas in the core of a jet engine?
Coming from a slightly related question, I want to know what is the speed of the gas in the core of a jet engine, after combustion but before the first turbine stage.
I'll just pick a random example: Rolls Royce Trent XWB.
Also, I don't know if the speed in the core is equal to the speed at the exhaust. I just want to be clear that it's the core air speed I'm interested in here.
I realize core air speed might vary with thrust and flight level. In that case I think the most sensible thing to do is choose some typical flight conditions during cruise.

Would this answer your question?

@PeterKämpf It's a very good chart but it doesn't say what engine it's for, and also what operating conditions (cruise or static on the ground). From the looks of it, it might be an old turbojet. I don't know if all turbo engines have the same or very similar core conditions for all 3 parameters, especially if talking over the whole history of jets.

@PeterKämpf I just read all 13 pages of that pdf, "3_engine_fundamentals". It's very well written, but I can't find an author, publisher, or date of publication anywhere. Is it an excerpt? If so, what and where is the full text? BTW page 6, The overall pressure ratio ... is approximately 9.5:1. That makes it sound very dated. Modern big jet engines get 20:1 compression ratio or more, and pressure ratio will be greater according to P1/P2 = (V1/V2)^gamma.

In addition, i think i found an error on page 12. centrifugal flow compressors operate at somewhat lower pressure ratios than axial flow compressors. The rule of thumb I've heard here on this site is that centrifugals typically have 4:1 compression ratio, but axial stages have more like 1.2:1. And those are compression ratios, so pressure ratio of centrifugals will be even greater according to that same equation.

All your comments are correct. As far as I know, the diagram is made by RollsRoyce and is not for a specific engine, but shows generic data maybe from the Seventies. Regarding centrifugal compressors: Per stage they have much higher ratios than an axial compressor, but overall the compression ratios of jet engines with radial compressors are much lower than those with axial compressors. I think that is what the PDF means.


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Does Betz's Law apply to jet engine turbines?
Betz's Law says the extractable power from wind is only 59% (16/27) of its original power.
$$P_{max} = \frac{16}{27} \frac{1}{2} \rho A v^3$$
Rho is the gas density, A is the cross section area of gas flow, and v is the original flow velocity (before it goes through the turbine stages).
Note that power is force times velocity, so the force component is $\frac{1}{2} \rho A v^2$ and then multiplied by another v makes it $v^3$. This is a little confusing because it looks very similar to the formula for kinetic energy ($\frac{1}{2} m v^2$). However, Betz's Law deals with power.
Note: I'm asking about the gas flow after it has been combusted. So it starts out hot and fast with some speed and temperature. And we assume the total heat energy remains constant (it's an adiabatic process) because again, this is considering the gas after combustion.
So does Betz's Law apply to it? If not, what are the exceptions invalidating it? I'm interested in any turbine-based engine (turboprop, turbojet, turbofan).

Did you read the Assumptions section of the Wikipedia article? Particularly point 3 does not seem to hold for jet engines: "The flow is non-compressible. Density remains constant [...]"

@Bianfable Yes i read those 4 assumptions, and point 3 is very confusing to me. Air is very, very compressible. Yet Betz's Law applies to wind turbines, and what is wind? Moving air. So i find that very strange.

Betz Law does not have anything to do with gas turbines, it is about extracting kinetic energy purely from the wind itself, without any additional sources of energy (i.e. fuel).

For airflow < 100 m/s it can be treated as incompressible, which seems fine for wind turbines, but not for gas turbines.

@JuanJimenez The gas turbine extracts energy from air. This air is at higher temps, density, and pressure, but as far as i can tell the working principle is the same. Those turbine blades are airfoils and you deal with lift and drag just like the blade of a traditional wind turbine.

@OrganicMarble I should prolly ask another q about how fast is the air in a core of a turbine engine.

@DrZ214, see https://aviation.stackexchange.com/q/29279/524: about 300–400 m/s throughout the turbine, taking energy more from the pressure and temperature than the velocity.

@DrZ214 No, your understanding of how a gas turbine works is totally incorrect. You are confusing the combustion process with the generation of thrust. A gas turbine does not extract "energy from air," it gets its energy from the combustion of fuel with air. A wind turbine, which is what Betz Law addresses, gets its energy purely from the movement of air against the blades.

@JuanJimenez Someone deleted my comment. (EDIT: oops, wrong post, it's in Sanchises answer). I wouldn't say that. The point of a turbine is to extract power from a fluid (which either decreases its velocity or makes it expand). The point of jet exhaust is to increase the velocity of a fluid to make thrust. So they are kinda at odds with each other. I think you are confusing the different stages of an entire engine and trying to think of it as one big thing. My question is only dealing with the turbine. A turbine extracts energy from a fluid. A compressor applies energy to a fluid.

@drz214 your statement still makes little sense. I suggest you look up "gas turbine" and "wind turbine" on Wikipedia. The gas turbine gets its energy from the combustion, not "wind."

Betz limit applies to open systems, not to ducts. It is relevant for water too, but tidal turbines using channels are able to extract a lot more energy than the Betz allows, see Exceeding the Betz Limit with Tidal Turbines. Turbine engines have ducted paths too.


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What is the daytime visibility from within a cloud?
I ask this question from the context of formation-flying, such as during WW2. What is the visibility in a typical cloud during daytime? Can you lose your wingman right away? Were you at serious risk of mid-air collision when this happened?

I don't think you can really specify a "typical" cloud, but you could easily compare it to the visibility in various degrees of fog.  Or if you've flown on a commercial airliner, you may have noticed that the wingtips are usually easily visible when flying in clouds.

The answer is that visibility varies depending on the density of the cloud.  And it can vary significantly.
Flying in formation in IMC is very common in the military.  Sometimes it will be dense enough that all you will have is a wingtip or light.  In those cases it can be very disorienting and difficult to maintain position because you need more than a single point to discern relative motion.  The stress level can go up too since the denser the cloud the closer want you fly to keep sight.  But generally those sort of conditions are relatively brief, and within 10-15 seconds you will likely have the wing and fuselage back in sight.
For reference, in close formation the wing tip might be 10-20 feet from your canopy, with the lead's fuselage another 20 feet beyond that.  So, the worst visibility I have seen (when I lost sight) could have been anywhere from 0-20 feet.  If you see the lead clearly the visibility could be anywhere 50 feet to a mile or more.  You just don't know, all you see is another airplane and gray...  There is certainly no absolute or standard number.
Every formation flight brief should cover lost sight procedures in case it becomes dense enough to require a break away.  I have only ever lost sight and had to break away once.

ditto to your comments Michael.  Only thing I might add is that night-time makes it 10x worse. One of worst memories I have is on the boom of KC-135 at 1 or 2 am in the morning, at 20k MSL or so, in and out of very dense thunderstorm clouds over northern Thailand. After a few minutes I suddenly noticed my upper thighs were really aching, and realized I was pushing on the rudder pedals so hard I had lifted my ass off the seat by a half inch or so.... Intermittent lightning didn't help either...

I read an article a while back about how P-51s did arrivals back to base in IMC (after finding the base by ADF) after a mission, and it was pretty hair raising. They would open the formation for the letdown, although the leaders and wingmen would stay tight to each other to keep visual contact, and depended on very precise heading and speed control to keep separation until they broke out at the bottom.  Midairs were pretty common. Even more hair raising is the British night bomber streams, which on a moonless night depended totally on a relatively wide distribution, and chance, for separation.


text:
How do jet engines cool the gas in the combustion zone?
The adiabatic flame temperature of kerosene in air is 2093 C (source). Modern jet engines have a turbine inlet temperature (TIT) of 1700 to 1800 C. This is the max acceptable temperature before the turbine blades melt or weaken too much.
So how is the gas cooled to the lower temperature before it reaches the turbine stage?
Now I know the obvious simple answer is "air cooling". But I want to know some details about how it's achieved.

Is the cooling air coming from the core air, or is it coming from some ducts into the combustion zone?

Is the cooling air pumped by something (other than the compressor stages) near the core?

Is the cooling air redirected turbulantly by some geometry inside the combustion zone?

Is the cooling air expanded to achieve adiabatic cooling (again I'm only asking about inside the combustion zone)?

Is fuel circulated around the core to act as a coolant (like rocket engines)?

When searching this, I came across something called inlet air cooling. This is not what I'm asking about. Inlet air cooling is about cooling the air before it enters the engine, for example by evaporative cooling.
Also, there is technology to cool the turbine blades directly, such as drilling holes in them and passing some other air . This is not what I'm asking about either. I want to know how combustion gas is cooled before it reaches the first turbine stage.

I realize of course that I can't get a full teaching of fluid dynamics from any answer on this site. What I'm hoping for at the very least is pictures, diagrams, and preferably some numbers used in certain equations that relate temperature and mass flow of air.
One more thing. Old engines are okay. In fact, they might be better, because older engines have a lower TIT and therefore must achieve even more cooling in the combustion zone. I did try to research them, e.g. the Rolls Royce Nene here, where a paragraph about cooling is summarized in a final sentence: "This diluting air is consequently expanded and accelerated rearwards, but it also cools the products of combustion to the temperature required at the turbine inlet." But it has very few numbers and no diagrams or cutouts.

Blades and vanes can not tolerate temperatures anywhere near 1700 C.

There's the fact that the hot gases from combustion are actually cooled in the combustion chamber. The turbine cooling isn't sufficient enough to maintain the appropriate temperature for the turbine blade(~1700 C).

The above figure(shamelessly taken from Aircraft Propulsion) clearly depicts the cooling mechanism in the combustion chamber(CC). 12% of air coming out of the compressor is primarily used for combustion and the rest of the air is used for cooling the hot gases. The temperature of the primary zone where the fuel is ignited can be greater than 2000 C. The cooling is done through film, transpiration, and impingement cooling. 20% of the air in the primary zone is used for better combustion efficiency and flame stability while (40%+20%=60%) of air in the dilution zone is used to maintain desired Turbine Inlet Temperature.
Now back to your question:

Is the cooling air coming from the core air, or is it coming from some ducts into the combustion zone?

The core air is only used for cooling.

Is the cooling air pumped by something (other than the compressor stages) near the core?

The cooling air is only pumped by the compressor.

Is the cooling air redirected turbulently by some geometry inside the combustion zone?

The above figure clearly answers this question.

Is the cooling air expanded to achieve adiabatic cooling (again I'm only asking about inside the combustion zone)?

As already mentioned, the cooling is done by transpiration, film, and impingement cooling.

Is fuel circulated around the core to act as a coolant (like rocket engines)?

No, the fuel isn't used for cooling like in a rocket engine. It's inefficient as fuel isn't stored at cryogenic temperature.

Your source about the flame temperature is valid for a stoichiometric mixture. This means all oxygen molecules will be matched by carbon and hydrogen such that theoretically a complete conversion into H₂O and CO₂ can take place.
Real gas turbines, however, operate with a fairly large amount of excess air, which is why they produce comparably large amounts of nitrous oxides. This also means that energy from the combustion is not only heating the molecules involved in the combustion process, but also further oxygen molecules which will either remain or react with nitrogen. This excess oxygen is what allows afterburners to function - if all oxygen would be used up in the combustion chamber already, none would be left for the afterburner.
This excess air is used for dilution and as a cooling agent, and by being forced through holes or a porous structure in the single combustion cans or the liners of an annular combustor cools the structure defining the combustion chamber. To cite Wikipedia:

… combustors are carefully designed to first mix and ignite the air and fuel, and then mix in more air to complete the combustion process.

and further down, in the section about the liner:

… even though high performance alloys are used, the liners must be cooled with air flow.

The gas is not cooled prior to turbine inlet entry.   Any attempt to cool down the gas, with decreasing efficiently of the engine.  Most modern engines have turbine inlet temperatures of approximately 3600° F.  This is far too hot for conventional alloys used to make the stator of rotor buckets to maintain acceptable yield strengths, so engineers have been forced to use the following techniques for manufacturing hot section components.

Manufacturing parts from exotic ceramic materials or ceramic hybrids which offer the required material properties at these temperatures.

Manufacturing parts form exotic metals such as nickel, cobalt, molybdenum, titanium or alloys of these.

Exotic single crystal casting techniques which offer superior material properties to conventional manufacturing processes.

Bleed air boundary layers - bleed air is tapped from the HPC and exits turbine casing, stators and rotors through small, laser drilled holes in the surface of the parts, effectively enveloping the part in a boundary layer of cooler air, preventing hot gas from making contact with the blades.

The gas is not cooled prior to turbine inlet entry. This doesn't make sense to me. If the flame temp is 2093 C and the TIT is 1800 C, then obviously it was cooled. Any attempt to cool down the gas, with decreasing efficiently of the engine. Indeed, cooling the gas does decrease efficiency because we are creating a smaller difference in temperature in the Brayton Cycle. But it's necessary. The turbine can only handle 1800 C so there is no choice, we must cool it to 1800 C.


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Are turbine blisks common in jet engines?
This comes from a slightly related question about blisks. A blisk is a single component shaped like a disc with blades on it. This can be for compressors or turbines.
I want to ask about the turbine possibility here. Are turbine blisks common on relatively new jet engines, either commericial or military? It's arbitrary, but take "relatively new" to mean designs that went into production within the last 10 years, maybe give or take a few years.
The turbine has the most difficult alloy, superalloys, so I am wondering if blisk manufacture is made more difficult because of this or is infeasible for some reason. (Note the blisk has several disadvantages, like if one blade is damaged the entire blisk needs to be taken out and repaired/replaced, but that alone doesn't suggest to me why turbines would be more problematic than compressors.)
Just for reference, the first turbine stage needs to handle temperatures of 1700 to 1800 K.

To my knowledge, there are no turbine blisks in any current engine.  There may be one out there I don't know about, but they certainly are not common.
The reason is blade life.  With a blisk, if you have to replace one blade, you have to replace the entire blisk.  Repairs due exist, of course, but if it is too badly damaged to be repaired you have to replace everything. With separate blades, you can replace only the ones that need to be replaced.
Turbine blade life is much shorter than compressor blade life, due to the extreme temperature environment.  A typical engine might go through multiple entire sets of turbine blades in its lifetime.  An average compressor blade might last the entire lifetime of the engine.

You answer was nearly 5 hours after the question.  Next time, be more blisk about it.  If you can touch type like me, you can type very bliskly, improving matters further.

@JohnK That's just bliskful thinking.

I'm in aircraft engineering school right now, and i can tell you blisks are definitely in use.
For example the Rolls-Royce Trent 500 turbofan, or the Allison Model 250.
(The latter of which we have a cut-open version of, on a dolly here)
This is the turbine section as blisk (edit: corrected from fan)

And this is the compressor stages as blisk vvv

If you have any questions, i'll be glad to elaboate. :)
Edit: I saw you wanted more recent engines!
The Hf-120 turbofan, first run in 2009 utilizes blisk technology

Edit 2:
Even more recent is the R-R Trent XWB, with new blisk drums. Certified in 2013

Oberusel factory reached a milestone of shipping out 10.000 blisks in 2019, for this engine
(europawire.eu)

Oh, also the The Eurojet EJ200 engine. And I just realised you wanted an answer about engines within 10 years ;_;

Thank you but my OP explicitly says turbine blisks, not compressor blisks or fan blisks.

Sorry, that's a mis-definition by me, the first picture is a blisk turbine (on the Trent 500)

Give me a little time, i'll pass the question by my professors